The present invention is applicable to a class of rocket engines that provide vacuum thrust covering a range of approximately 500 lbf-20,000 lbf and 25,000 lbf-100,000 lbf, featuring advanced swirl combustion, and can be used to meet the throttling propulsion needs for traveling to the Moon, Mars and beyond.
The liquid propulsion engine upper stage and spacecraft rocket engine market, continues to receive emphasis, especially for storable non-toxic and cryogenic propellant systems, high expansion nozzles and orbit maneuvering systems. For current upper stage Expendable Launch Vehicles (ELV), the generated vacuum thrust varies from approximately 25,000 lbf-60,000 lbf for Atlas 5 and the Delta-IV (small, medium and heavy) launch vehicles. Future potential growth is anticipated in this mid-thrust range liquid ELV rocket propulsion market, with expected vacuum thrust needed in the range of approximately 25,000 lbf to about 100,000 lbf.
The recent published NASA Exploration Systems Architecture Study (ESAS) reviewed numerous propulsion options to determine how the Crew Exploration Vehicle (CEV) and the Crew Launch Vehicle (CLV) could be utilized to transport both crew and cargo to the International Space Station (ISS), as well as transportation of crew and cargo to the Moon and Mars. The ESAS concluded that a variety of propulsion technologies are required to be developed in order to support missions to the ISS, returning to the Moon, and future missions to Mars. There were three new propulsion project recommendations from the ESAS architecture to support the above missions: (1) human-rated, 5,000 lbf-20,000 lbf class in-space propulsion engines to support the Service Module (SM) for ISS orbital operations, lunar ascent and Trans-Earth Injection; (2) human-rated deep throttleable 5,000 lbf-20,000 lbf thrust class engines for lunar descent; and (3) human-rated pressure/pump-fed 5,000 lbf-20,000 lbf thrust class engines for the upgraded Lunar Surface Access Module (LSAM) ascent stage.
The above architecture study specified a high specific impulse (Isp) propulsion system for the SM and lunar ascent that would provide high reliability without significant propellant boil-off. A human rated 5,000 lbf-20,000 lbf pressure/pump fed Liquid Oxygen (LOX)/Methane (CH4) in-space propulsion engine has been specified in the study for both the SM and LSAM ascent stage and for the upgraded pump-fed version.
For the LSAM descent stage, a throttleable 5,000 lbf-20,000 lbf pressure/pump-fed (gas generator or expander turbo-pump feed system), deep-throttling engine has been identified to support the return to the Moon. The propellants chosen consist of LOX/liquid hydrogen (LH2). The LOX/LH2 pump-fed propulsion engine was selected for the lunar descent stage due to its higher Isp performance, lower cost and risk level relative to a pressure-fed system, which will allow the LSAM to perform a circular burn, yet maximize the LSAM cargo delivery capability. Furthermore, these engines must have a restart capability for lunar descent and the ability to throttle down to 10 percent of the total thrust. Conversely, common pressure-fed LOX/CH4 engines were chosen for the CEV Service Module and lunar ascent stage propulsion systems in order to attain high reliability performance engines with similar propellants.
Thus, there is a need to develop versatile rocket propulsion engines to support NASA's near-term propulsion requirements with vacuum thrust throttling capability at least covering the 5,000 lbf-20,000 lbf range using LOX/LH2 and LOX/CH4 for orbital maneuvering, Lunar and Mars descent, landing and ascent, as well as a complimentary vacuum thrust ranging up to approximately 60,000 lbf and beyond to address Trans-Earth Orbital Insertion from Mars and Trans-Lunar/Mars Insertion, Mars/Lunar surface hopping, as well as for other applications. Additionally, mid-thrust propulsion engines in the 25,000 lbf-100,000 lbf vacuum thrust range are expected to be required to support the future ELV upper stage market.
Compact Advanced Swirl-combustion Propulsion (CASP) technology that can be used in a broad range of rocket-based and air-breathing propulsion applications has recently been successfully developed and test demonstrated. It is simple and has no moving parts. The swirl generator is the central key feature of the CASP technology, which enables robust mixing of propellants, flame stabilization and flame propagation that produces near complete combustion over a wide range of compact combustor lengths and diameters.
CASP technology was originally developed and tested as an auxiliary ramjet thrust propulsion system to provide additional lift thrust augmentation for the Boeing Joint Strike Fighter (JSF) Short Take-off Vertical Landing aircraft (See “COMPACT, LIGHTWEIGHT HIGH-PERFORMANCE LIFT THRUSTER INCORPORATING SWIRL-AUGMENTED OXIDIZER/FUEL INJECTION, MIXING AND COMBUSTION,” U.S. Pat. No. 6,820,411 by Pederson et al., which is incorporated by this reference.) Using jet propulsion fuels and simulating air off-take from the JSF gas turbine engine fan, the ground-tested swirl ramjet engine delivered very high engine performance margins (nearly ideal) that consisted of a measured gross fuel specific impulse of 3300 to 2100 seconds (ideal fuel Isp is 3500 to 2300 seconds), with corresponding combustion efficiencies of 99% to 90% over an equivalence ratio (fuel/air ratio) range of 0.5 to 1.0. Propulsive efficiency of the CASP engine was approximately constant at 93% over mid to high equivalence ratio ranges (growth capability to about 96%) and consistently higher than historical ramjet Lightweight engines, yet the CASP combustor length was 36% shorter. Test-to-test performance repeatability of the CASP was excellent, with excursions below ±2.2%. The tested combustor/convergent nozzle length to combustor diameter ratio, L/D, was only 1.6, with further potential for reducing the L/D down to 1.0 or less. In an airbreathing ramjet configuration, the CASP technology was tested with extreme length constraints and at very difficult operating flight conditions compared to typical ramjets, which can adversely affect fuel injection, ignition and combustion stability, yet these were overcome and 5:1 engine throttleability was test demonstrated. The CASP technology also demonstrated smooth combustion (high frequency pressure fluctuations<5%), which underscores the viability and practicality of this developed technology.
The swirl generator design is quite flexible and has been implemented in many other propulsion applications with similar benefits. For example, it has been employed to introduce novel: (1) Compact, Lightweight Ramjet Engines Incorporating Swirl Augmented Combustion With Improved Performance, (see U.S. Pat. Nos. 6,968,695 and 7,137,255 by S. Schmotolocha et al.); (2) Compact Swirl Augmented Afterburners for Gas Turbine Engines (see U.S. Pat. Nos. 6,895,756 and 7,137,255 by Schmotolocha et al.); and (3) Combined Cycle Engines Incorporating Swirl Augmented Combustion for Reduced Volume and Weight and Improved Performance (see U.S. Pat. No. 6,907,724 by R. Edelman et al.).
This propulsion technology that was originally developed for airbreathing applications could also be applied to liquid rocket engines, which could lead to the next generation of ELV's and support NASA's propulsion requirements for the return to the Moon, as well as future Mars missions. There is great potential for transferring this advanced swirl propulsion technology from airbreathing propulsion engines to rocket engines. The resulting key benefits to NASA, DoD and the commercial launch market for small-to-medium size rocket engines (500 to 100,000 lbf thrust range) would include: (1) a significant reduction in combustor length, weight and complexity; (2) a greater nozzle expansion ratios to provide higher vacuum Isp can be introduced into fixed rocket engine lengths due to shorter combustor length; (3) reduced combustor cooling requirements due to a shorter combustor length; (4) improved engine performance (high Isp) and operability; (5) engine throttling ability; and (6) enhanced safety and lower system part count, which results in reduced complexity and manufacturing cost. The CASP propulsion system is highly reliable and the attendant propellants injection approach is extremely flexible. The latter allows the engine design to be tuned to achieve near maximum theoretical performance (thrust, Isp, propulsive and combustion efficiencies).
These advantages would not be limited to a specific rocket engine type, but can be applied to a wide variety of medium-to-small rocket engine sizes. Typical feed systems used to fuel rocket propulsion combustors, such as simple gas pressure or turbo-pump feed systems, have limited throttling ratios (the rocket engine's maximum thrust capability compared to the minimum thrust capability). Gas pressure feed systems are capable of achieving throttling ratios of approximately 10:1, such as the pressure-fed Apollo descent engine. Turbo-pump feed systems currently achieve throttling ratios of 5:1. Additionally, a rocket engine that utilizes a gas pressure feed system operating at low tank pressure and low combustion chamber pressure (Pc) results in an undesirably large sized rocket engine that has poor propulsion performance. Conversely, if the same rocket engine is required to operate at high Pc and high tank pressure, the weight of tank increases to an undesirable size and weight. Turbo-pump feed systems allow rocket engines to operate at a high Pc and a low tank pressure, and simultaneously provide higher propulsion performance than a simple gas pressure feed system. This rocket engine is, however, significantly more complex and, as with the gas pressure feed system, is unproven for use with CASP technology. Thus, there is a need for the swirl combustion technology to be incorporated into small to medium size rocket engines to provide high throttling capability for either a gas pressure or a turbo-pump feed system that uses the fuel in a regenerative cooling approach to cool the MCC and nozzle, resulting in smaller and lighter vehicles with high performance. This new type of rocket engine design would be a significant improvement over traditional rocket engine design for space applications.